1. Field of the Invention
The present invention relates to systems and methods for controlling spacecraft, and in particular to a system and method for controlling the spacecraft while deploying a large appendage such as a reflector.
2. Description of the Related Art
Satellite systems typically include appendages that are stowed during launch and deployed sometime thereafter. Typically, these appendages include solar wings (having the solar cells and the supporting structures), passive and/or active sensors, and antennas used for satellite-to-ground or satellite-to-satellite communications.
A typical goal during satellite appendage deployment is to maintain a spacecraft dynamic state close to that of rotating about an axis fixed in both the body and in inertial space. A rotation axis fixed in both the spacecraft and inertial space implies that the satellite angular rate vector direction is fixed in the satellite. This is an idealized state, and deployments are designed to keep the spacecraft state within acceptable bounds of this state.
One purpose of this goal is to keep the sun close to the desirable direction for the body, for solar power or thermal control reasons. Another is to keep the axis of symmetry of the toroidal telemetry and command antenna close to a known direction in space. The rotation itself helps largely to average out momentum buildup due to environmental torques, and ensures that any blockages of telemetry and command antenna line of sight by spacecraft structures are cleared by rotation. Such appendages must also be deployed in such a way that they are not damaged, and such that the long term stability of the satellite is not adversely affected.
Fortunately, since the inertia of each of the appendages is typically much smaller than that of the satellite itself, the attitude control system of the host satellite is typically adequate to assure the stability of the satellite during and/or after deployment.
A number of solar array deployment techniques have been developed. A first technique uses only the design variables of the satellite spin speed and the initial orientation of the spin vector in space, and the spacecraft North and South solar arrays are deployed in turn without any other attitude control. A second technique times the first deployment with respect to the sun rotational phase, and the second deployment with respect to the first, in order to meet the design constraints. A third technique initiates nutation with a thruster pulse or a wheel maneuver prior to the release of the first solar array, timed so that the constraints of the deployment are met over the deployment sequence. A fourth technique uses an internal wheel momentum to alter the dynamic motions during deployment to favorable effect. This includes using an internal wheel momentum bias along the spin axis (“superspin”) to make the effective inertia ratio σeff>1, thus rendering the spacecraft passively stable during the deployment. A fifth technique uses active wheel control during deployment to help damp out unwanted spacecraft body rates. Other deployment techniques have used similar methods, even for deployment of asymmetrical antennas measuring 49 feet in maximum dimension, and over long duration deployments lasting several hours.
In all of these techniques, the effects of momentum buildup due to environmental torques (including solar torques, thermal emission torques, atmospheric drag torques, and radio frequency (RF) emission torques) are small enough so that the effect on the satellite attitude is acceptably small, and are therefore neglected.
However, in some circumstances, the effect of environmental torques and the momentum of the deployed appendage(s) cannot be adequately accommodated by the attitude control system of the satellite. This can be the case, for example, when the satellite appendage is a large antenna.
Satellites have long been used to communicate information with terrestrially based ground stations. Communications systems have also been devised to permit worldwide communications to mobile receivers, but such designs have met with limited success. One of the reasons for this limited success is that it is difficult to design a lightweight, mobile receiver that can communicate with the satellite constellation, even a constellation in mid or low Earth orbits.
One solution to this problem is to equip one or more of the satellites with one or more high-sensitivity transmit/receive antennas. The problem with this solution is that to provide high sensitivity over wide beamwidths, the antennas must be quite large. This raises difficulties in three respects. First, the larger antennas have larger inertia than the smaller prior art antennas discussed above, thus imposing more stringent requirements on the design of the satellite control system, including the sensors (e.g. star and sun sensors, gyros and accelerometers), the attitude control thrusters, and the algorithms implemented by the control system. Second, since the larger antenna must be designed to fit within the storage bay of the launch vehicle deploying the satellite, the structures relating to its deployment are typically more complex than those of smaller appendages. For example, a foldable large antenna will generally have more joints and more structural elements than a smaller prior art antenna, and each of these structural elements must be folded out during deployment. Expressing the dynamics of the deployment of such antennas can be a daunting task, making it difficult for the satellite designer to assure that such deployment does not compromise satellite stability.
It is also desirable to perform the deployment over an extended period of time to avoid mesh tangling, and to even the thermal condition of truss joints. However, doing so subjects the satellite and the deployed appendage to a variety of disturbance torques (including solar torques) that further complicate the deployment process. It is possible, of course, to simply design the attitude control system to permit the rapid deployment of the appendage (e.g. by thruster torques), but such a solution would require a satellite control system that would be far more robust than is required at any time after the appendage is deployed, increasing the weight and cost of the satellite. Also, using thrusters during deployment raises the possibility that the appendage will be damaged during deployment.
Solar torquing techniques using satellite appendages have been applied to operational non-spinning satellites with relatively constant mass properties. For example, such techniques are disclosed in “Survey of Solar Sailing Configurations for Satellite Attitude Control,” AAS 91-486, by George A. Kyroudis, which is hereby incorporated by reference herein. U.S. Pat. No. 5,816,540, issued to Murphy, U.S. Pat. No. 4,949,922, issued to Rosen, and U.S. Pat. No. 4,325,124, (also incorporated by reference herein) also disclose oscillating one or more of the solar arrays to apply solar torque (a technique sometimes known as “solar tacking”). While useful, such techniques are thought to be only applicable in operational, non-spinning satellites with relatively constant mass properties (e.g. not during deployment of a large appendage). Further, such techniques are limited in application, because the manipulation of the solar arrays can have a deleterious effect on the ability of the solar arrays to generate power for satellite operation.
Transverse momentum has been used to produce cross-product gyroscopic torques in other contexts, namely to oppose external boost motor torques. This technique is disclosed in U.S. Pat. No. 6,032,903, which is hereby incorporated by reference. However, this technique relies on a combination of gyroscopic torque and thruster torque, and, as described above, it is undesirable to use thruster torque during appendage deployment. Further, this technique is typically appropriate for different and much larger disturbance torques.
Magnetic torquing momentum management systems have also been used. Such techniques use a small gimbal offset in a gimbaled momentum wheel, such that, as the satellite body rotates 360 degrees in pitch over 24 hours, the peak transverse angular momentum over the day is largely minimized. However, this technique is not generally applicable to the problem of large appendage deployment, where the dynamic state is expected to be a rotation perpendicular to the pitch axis, at a rotational rate that is typically an order of magnitude higher.
What is needed is a system and method for controlling a satellite while deploying an appendage having a large moment of inertia. What is also needed is that the method be sufficiently robust to permit the design of a deployment sequence without requiring a detailed model of the inertia of the appendage as a function of time during the deployment sequence. The present invention satisfies that need.